Small gas turbine engine with unmixed exhaust

ABSTRACT

A small twin spool gas turbine engine with a bearing support arrangement in which the bearings are dampened by O-rings secured between the bearing races and the support structure, the bearings are arranged in series so that a cooling air can be passed through the bearings to prevent overheating, and the bearings are dry lubricated. The aft end of the engine includes high speed and low speed ball bearings supported with preload springs to add additional damping capability. The end bearing support assemblies form a cooling air path for the cooling fluid to flow through a passage within one of the guide vanes, through the bearings and a hole in the low pressure turbine rotor disk and out the exhaust of the engine. A snorkel tube extends into the bypass air channel drawing air for the bearings, and has a slanted opening preventing dirt particulates from entering the bearing cooling passage.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is claims the benefit to an earlier filed U.S.Provisional application 60/963,084 filed on Aug. 2, 2007 and entitledSMALL TWIN SPOOL GAS TURBINE ENGINE the entire disclosure of which isherein incorporated by reference.

This application is also related to U.S. Regular application Ser. No.11/903,553 filed on Sep. 21, 2007 and entitled HIGH SPEED ROTOR SHAFTFOR A SMALL TWIN SPOOL GAS TURBINE ENGINE; and related to U.S. Regularutility application Ser. No. 11/903,555 filed on Sep. 21, 2007 andentitled HIGH SPEED ROTOR SHAFT AND TURBINE ROTOR DISK ASSEMBLY; andrelated to U.S. Regular utility application Ser. No. 11/903,554 filed onSep. 21, 2007 and entitled HIGH SPEED ROTOR SHAFT AND COMPRESSOR ROTORDISK ASSEMBLY; and related to U.S. Regular utility application Ser. No.11/975,674 filed on Oct. 19, 2007 and entitled LOW PRESSURE TURBINEROTOR DISK; all the above of which are incorporated herein by reference.

FEDERAL RESEARCH STATEMENT

The US Government has a paid-up license in this invention and the rightin limited circumstances to require the patent owner to license otherson reasonable terms as provided for by the terms of Contract No.FA9300-04-C-0008 awarded by the United States Army.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates generally to a small twin spool gasturbine engine, and more specifically the cooling of the rear endbearings.

2. Description of the Related Art Including Information Disclosed Under37 CFR 1.97 and 1.98

In a gas turbine engine, the rotor shaft or shafts is supported forrotation by at least two bearings which include a forward or compressorend bearing and an aft or turbine engine bearing. Both radial and axialloads must be absorbed by the bearings. It is typical in the larger gasturbine engines of the prior art to use a ball bearing in the compressorend and a roller bearing in the turbine engine. The ball bearing canabsorb both radial and axial loads and therefore acts as the thrustbearing for the rotor shaft. Because of the high temperatures that occurin the turbine end, the rear bearing must be capable of allowing axialdisplacement between the outer race support housing for the bearing andthe bearing itself. This is why roller bearings are used in the turbineend. However, roller bearings only provide radial load absorption and noaxial load. If a roller bearing is subject to an axial load, the rollerswill start to rotate and then wobble during rotation of the bearing.This is a very undesirable situation and usually will result in thebearing blowing itself apart.

In a small gas turbine engine of bellow around 300 pounds thrust, theturbine end of the engine is exposed to very hot temperatures. Theturbine end bearings are usually cooled by an external supplied coolingfluid such as a wet lubricant. Using a wet lubricant to provide coolingfor the bearing requires a wet lubricant reservoir and the pumping anddelivery system to circulate the liquid lubricant. This takes upadditional space and adds weight to the engine. For a small gas turbineengine, this could add double the size and weight of the overall engine.

A typical gas turbine engine, especially of the small size as in thepresent invention, the rear end bearings would be cooled by sealingthese bearings in a compartment that is filled with a wet or liquidlubricant. Or, some small engines pass fuel through the bearings toproduce lubrication and cooling, and then discharge the unburned fuelout from the engine. In these engines, a lubricant is required to bestored in the engine until the engine is used. For a missile that usesthis type of engine, long time storage is a problem because the fuel orlubricant cannot be held for long before it degrades.

BRIEF SUMMARY OF THE INVENTION

It is an object of the present invention to provide for a small twinspool gas turbine engine less than 300 pounds of thrust.

It is another object of the present invention to provide for a small gasturbine engine with a cooled aft end bearing that does not require a wetlubricant.

It is another object of the present invention to provide for a small gasturbine engine that can be stored for long period of time before beingoperated.

It is another object of the present invention to provide for a small gasturbine engine with an air cooled aft end bearing that is free fromparticulate matter that could damage the bearing.

It is another object of the present invention to provide for a small gasturbine engine that uses cooling air from the bypass channel to cool thebearings without the need of a wet lubricant.

The present invention is a small gas turbine engine having twin spoolsor rotor shaft in which the aft end bearings are ball bearings that arecooled by passing cooling air through the bearings in which the coolingair is bled off from the bypass channel through a snorkel tube thatextends from the bearing cooling channels and into the bypass channel todraw in the cooling air discharged from the fan compressor. The snorkeltube is open on the top end and angled to face the downstream directionof the bypass air path such that any particulate material such as dirtparticles flowing in the bypass air will flow over and past the snorkeltube opening. Thus, only clean air will flow into the snorkel tube to besupplied as cooling air for the aft end bearings. The air diverted fromthe bypass channel is passed through the guide vane and into the chamberin which the bearings operate. The cooling air passes through thebearings and is then discharged from the engine through holes formed inthe low pressure turbine rotor disk and into the exhaust at near ambientpressure.

The bearings also include race surfaces and balls coated with a drylubricant and a hard coating to reduce friction that produces heat andto allow for high speed operation.

The aft end bearings are also ball bearings that are mounted within thehousing to allow for the outer races to slide axially in order toprevent axial loads from building up on the aft end bearings. Thebearings are supported on the outer race by a O-ring that providesdamping to the bearings. The outer race is also biased in the axialdirection by a preload coil spring to prevent the balls from beingunloaded during all phases of engine operation.

Together, the aft end bearing structure and arrangement, along with thecooling capabilities, allows for the small twin spool gas turbine engineto be capable of operating at the high rotational speeds in order tomake such a small engine possible.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

FIG. 1 shows a cross section view of the aft end bearing assembly forthe small twin spool gas turbine engine of the present invention.

FIG. 2 shows a cross section view of the aft end turbine rotor disk andbearing assembly of the present invention.

FIG. 3 shows a cross section detailed view of the aft end bearingassembly of the present invention.

FIG. 4 shows a cross section view of the twin spool gas turbine engineof the present invention.

FIG. 5 shows a cross section view of the rear end bearing coolingcircuit with the exhaust diffuser.

DETAILED DESCRIPTION OF THE INVENTION

The present invention is a bearing assembly and arrangement for a smalltwin spool gas turbine engine that allows for such a small engine toovercome the problems with rotor dynamics and cooling of the bearings.FIG. 1 shows the bearing assembly for the aft end of the engine. Theinner or low speed rotor shaft 11 is supported by an aft high speedshaft bearing 12 and the outer or high speed rotor shaft 13 is supportedby the aft low speed shaft bearing 14. The two aft end bearings 12 and14 are arranged in series such that their rotational axis is aboutaligned with each other. Each bearing 12 and 14 includes an inner raceand an outer race with roller balls secured between the races. The outerrace is thicker in order to reduce hoop stresses.

The aft end bearings 12 and 14 are supported by the low pressure turbineguide vane assembly 21 which includes a guide vane 22 extending inwardtoward the bearings 12 and 14, an inner guide vane shroud 23 and abearing support surface 24. The guide vane assembly that supports theaft end bearings 12 and 14 is formed of two pieces 23 and 24 that arebrazed together to form a single rigid piece. The inner surface of theaft bearing support surface 24 forms an annular surface on which the twoaft bearings are rotatably supported for operation in the engine. A seal29 is formed on the inner surface of the guide vane assembly 21 to forma seal with the tip of the first stage rotor blade in the turbine. Axialholes are formed to secure the guide vane assembly 21 to the rest of theengine as seen in FIG. 5.

The inner bearing support surface also includes two annular grooves thatopen inward toward the rotor shafts in which an O-ring 26 is supported.The O-rings provide for the required damping of the bearings in order toallow for the high speed rotation of the engine without exceeding therotor dynamics issues. The O-rings are formed of a high temperatureelastomeric material such as Parker FF200 which is capable ofwithstanding a temperature of up to around 615 degrees F. The O-ringshave a major diameter of 1.082 inches and a cross sectional diameter of0.070 inches. The bearings have a thicker outer race than on the innerrace in order to provide better hoop surface and to allow for the use ona single O-ring for damping. The O-ring can also be made from a materialreferred to as Calrez or Cham Raz.

In this embodiment, only one O-ring is used for each bearing so that theload applied to the bearing is not too high. With one O-ring perbearing, the load applied to the bearing race due to the O-ring may notbe centered properly. However, in another embodiment two O-rings areused for each bearing in order to more properly center the load appliedto the bearing race. However, the load would be higher using two O-ringsinstead of the single O-ring.

The balls of the ball bearings and the inner surfaces of the races arecoated with a lubricant and a hardener to provide for dry lubrication.The balls are formed of silicone nitride and coated with tungstendisulfide which is a solid lubricant. The inner races of the bearingsare formed from Rex 20 (could be stainless steel) and are coated withlayers formed from chromium nitride or titanium nitride (both hardcoatings) alternating with tungsten disulfide (a solid lubricant). Thus,alternating coatings of chromium nitride and tungsten disulfide ortitanium nitride and tungsten disulfide are formed on the inner and theouter races of the bearings. The bearings in the present invention haveno cages in order to eliminate the need for lubrication. Bearings withcages require lubrication. The coatings described above provide for abearing for use in the engine that does not require a lubricant. This isanother main feature of the invention that allows for the small twinspool gas turbine engine to be operational, especially for a one timeuse like in a cruise missile that has a flying range of around one hour.

A pre-load spring 27 is also used to secure the two aft bearings 12 and14 in place on the aft bearing support surface 24. In this embodiment,the pre-load spring 27 is a coil spring to limit the number of pieces. Asnap ring 28 that fits within a snap ring annular groove on the bearingsupport surface 24 secures the pre-load springs 27 in place when thebearings are secured. The pre-load springs 27 also provide for therequired damping of the bearings to allow for the high speed rotationand the small twin spool gas turbine engine possible.

The aft end high speed bearing 14 is secured to the high speed rotorshaft through the high pressure turbine rotor disk 31 by a nut 32 fromthe aft end of the bearing as seen in FIG. 2. An abutment surface isformed on the high pressure turbine rotor disk 31 on which the forwardend of the bearing 14 abuts to secure the bearing 14 in place when thenut 32 is tightened. The high pressure turbine rotor disk 31 includes anaxially extending portion 33 on which the bearing 14 is supported. Alabyrinth seal member 35 is secured on the low speed shaft 11 andincludes lab seal teeth extending outward to form a lab seal with theaxial extending portion of the high pressure turbine rotor disk 31.

The aft end low speed bearing 12 is secured in place between thelabyrinth seal member 35 and the low pressure turbine rotor disk 36 asseen in FIG. 2. Abutment faces are formed on these two members to securethe bearing 12 in place. A nut threaded over the aft end of the lowspeed shaft 11 is used to compress the low pressure turbine rotor disk36 onto the low speed shaft 11 and load the bearing 12. An anti-rotationpin 38 is secured within axial grooves formed on both the low speedshaft 11 and the rotor disk 36 to prevent relative rotation.

The aft end bearings 12 and 14 are cooled by passing cooling air bledoff from the fan compressor through the bearings and passages formedbetween the inner race and the housing and then out from the turbinerotor disk. The cooling passages are formed as axial extending groovesin the support structure of the inner race. In other embodiments, thecooling air passages for the outer races could be formed in the outerraces or in both the outer races and the housing surface abutting therace. The outer surface of the inner race encloses the axial grooves toform the passages. The guide vane 22 includes an inner cooling airsupply passage 41 to supply cooling air from a source such as bleed offair from the bypass passage of the engine.

A snorkel tube 95 fits into an opening of the outer shroud of the vaneguide assembly to provide for a cooling air connection between thebypass channel 96 and the supply passage 41 in the vane. As seen in FIG.5, the snorkel tube 95 extends into the bypass channel 96 and theopening is slanted away from the flow path direction of the air so thatany dirt particles will pass over the opening and not pass into thetube. The purge air passage 46 also connects to the bypass channel 96.The snorkel tube 95 extends out and into the bypass channel 96 in orderto prevent dirt particulates from entering the cooling air for thebearings. The purge air passage formed within the guide vane 22 ends atthe outer surface of the vane outer shroud as seen in FIGS. 1, 2, 5 and6.

The bearing cooling air passage is formed to divert cooling air from thebypass passage 96, pass the cooling air through the bearings 14 and 12,and then discharge the spent cooling air out from the engine. In orderfor the pressure differential to exist between the upstream end and thedownstream end of the bearing cooling passage, the cooling circuit mustbe separate from the exhaust gas passage of the turbine. The turbineexhaust diffuser is formed downstream from the guide vane 97 and extendsoutward from the engine at a longer distance than does the compressorbypass fan passage formed outward from it. The bearing cooling airexhaust passage is formed by the member 98 as seen in FIG. 6 and extendsfarther rearward than the turbine exhaust diffuser in order to dischargethe bearing cooling air at close to ambient pressure.

A sealed annular space 42 is formed between the two brazed pieces 23 and24 that form the guide vane and bearing support assembly which isconnected to the inner cooling air supply passage 41. Another coolingair passage 43 is formed in the bearing support surface 24 that opensinto an inner bearing space 44 in which the aft bearings 12 and 14 arelocated. One or more axial holes 45 are formed in the turbine rotor disk36 to allow for the cooling air and fuel mixture to exit the bearingcooling passage. Thus, to cool the bearings the source of air is bleedoff air diverted from the bypass air from the fan compressor. Thesnorkel tube 95 extends from the bypass channel and through the guidevane 22 where the cooling air passage opens into the space 42 formedbetween the guide vane and bearing support structure 24. The bearingcooling air then passes through the bearings 14 and 12 with some of thecooling air passing through the axial passages formed along the outerraces, and then is discharged out from the engine through the axialholes 45 formed in the low pressure turbine rotor disk 36.

As seen in FIG. 2, a purge air supply passage 46 is also formed withinthe guide vane 22 to supply compressed air to a rim cavity 48 formedbetween the guide vane 22 and the turbine rotor disk 36. The purge airsupply passage 46 is connected to the area where the high pressurecompressor discharges its compressed air which is outside of the outershrouds of the guide vanes. An opening or passage 47 is formed withinthe bearing support surface 24 to connect the purge air supply passage46 with the rim cavity 48. A labyrinth seal assembly is formed betweenthe guide vane inner shroud and the rotor disk 36 as is typical in theart of gas turbine engines. The purge air prevents the hot gas flow frominjecting into the rim cavity and adding heat to the bearings.

FIG. 3 shows a detailed cross section view of the aft bearing assemblywith the bearing support surface 24, the axial hole 47 for the purgeair, the radial hole 43 for the bearing cooling air mixture, the annulargrooves with the O-rings 26, labyrinth seal teeth 49 extending outwardthat form a seal between the static bearing support surface 24 and theinner surface of the rotating rotor disk 36, the two bearings 12 and 14,the coil springs 27, and the snap rings 28.

1. A twin spool gas turbine engine comprising: a low speed shaftsupported by a forward end bearing and an aft end bearing; a fancompressor rotatably supported on the low speed shaft; a high speedshaft supported by a forward end bearing and an aft end bearing; a guidevane assembly located between a high pressure turbine rotor disk and alow pressure turbine rotor disk, the guide vane assembly forming an aftend bearing support surface; and, a closed cooling circuit passingthrough the aft end bearings to provide cooling air for the bearings,the closed cooling circuit opening into a bypass channel of the engineand ending outside the engine exhaust at near ambient pressure.
 2. Thetwin spool gas turbine engine of claim 1, and further comprising: theclosed cooling circuit passing through the guide vane assembly.
 3. Thetwin spool gas turbine engine of claim 2, and further comprising: a lowpressure turbine rotor disk connected to the low speed shaft and havingat least one cooling air discharge hole to discharge the bearing coolingair.
 4. The twin spool gas turbine engine of claim 1, and furthercomprising: the aft end bearings for the low and high speed shafts arealigned in series at about a same radial distance from the enginerotational axis.
 5. The twin spool gas turbine engine of claim 1, andfurther comprising: a bearing cooling air exhaust channel formed by aninner wall of a turbine exhaust gas diffuser; the inner wall of theturbine exhaust gas diffuser extending further aft than an outer wall ofthe turbine exhaust gas diffuser such that the bearing cooling airexhaust is near ambient pressure.
 6. The twin spool gas turbine engineof claim 5, and further comprising: the outer wall of the turbineexhaust gas diffuser forms an inner wall of the bypass air channel; and,an outer wall of the bypass channel extends aft less than the distanceof the outer wall of the turbine exhaust gas diffuser.
 7. The twin spoolgas turbine engine of claim 2, and further comprising: a snorkel tubeextends from a guide vane assembly bearing cooling air passage and intothe bypass channel.
 8. The twin spool gas turbine engine of claim 7, andfurther comprising: the snorkel tube extends at around 90 degrees to aflow direction of the bypass channel, and the snorkel tube includes aslanted opening so limit the ingestion of dirt particulates into thebearing cooling air passage.
 9. The twin spool gas turbine engine ofclaim 2, and further comprising: the guide vane assembly includes apurge air passage extending from the bypass channel and opening into arim cavity formed between the guide vane assembly and the low pressureturbine rotor disk, the purge air passage providing purge air from thebypass channel into the rim cavity.
 10. The twin spool gas turbineengine of claim 1, and further comprising: the aft end bearings includecooling air passages formed between the inner race and the bearingsupport surface to allow for cooling air to flow around the inner races.11. The twin spool gas turbine engine of claim 3, and furthercomprising: the low pressure turbine rotor disk includes a plurality ofaxial bearing cooling air exhaust holes in an annular arrangement. 12.The twin spool gas turbine engine of claim 1, and further comprising:the aft end bearings are dry lubricated.
 13. The twin spool gas turbineengine of claim 12, and further comprising: the aft end bearings areball bearings.
 14. The twin spool gas turbine engine of claim 13, andfurther comprising: the aft end ball bearings are each supported by anO-ring on an outer race.
 15. The twin spool gas turbine engine of claim14, and further comprising: the aft end ball bearings each include theouter race preloaded by a spring.
 16. The twin spool gas turbine engineof claim 15, and further comprising: the preload springs bias the outerrace toward the front of the engine.